LABYRINTH SEALING ASSEMBLY FOR AN AIRCRAFT TURBOMACHINE

Abstract

A labyrinth seal assembly (40) for an aircraft turbomachine (10), comprising a rotor (42) surrounded by a stator (44), the stator (44) comprising an abradable element (48) surrounding the rotor (42) and defining an annular passage (P) around the rotor (42), the rotor (42) comprising rotor blades (46) extending radially outwards and passing through said passage (P) to limit the flow of an operating gas stream (F3) from upstream to downstream in said passage (P), characterized in that it further comprises, at said passage (P) and upstream of all the rotor blades (46), stator fins (50) projecting radially inwards from the abradable element (48) and profiled to axially straighten said gas stream (F3) by downstream of the stator fins (50). Figure for the abbreviation: Figure 2

Description

Title of the invention: LABYRINTH SEALING ASSEMBLY FOR AN AIRCRAFT TURBOMACHINE Technical field of the invention

[0001] The present invention relates to a labyrinth seal assembly for an aircraft turbomachine, as well as an aircraft turbomachine comprising such an assembly. Technical background

[0002] The technical background includes in particular the documents FR-A1-3 015 591, FR-Al-3 099 788, FR-A1-3 071 540 and FR-A1-3 001 759.

[0003] An aircraft turbomachine 10, such as that shown in [Fig.1], comprises a gas generator 12 which conventionally includes, from upstream to downstream, with reference to the flow of gases in the turbomachine 10, at least one compressor 14, 16, an annular combustion chamber 18 and at least one turbine 20, 22.

[0004] In the case of a twin-body turbojet, respectively low pressure and high pressure, the gas generator 12 comprises successively a low pressure compressor 14, a high pressure compressor 16, the combustion chamber 18, a high pressure turbine 20 and a low pressure turbine 22.

[0005] The gas generator 12 defines an annular flow channel V1 of a gas flow which passes through the compressors 14, 16, the combustion chamber 18 and the turbines 20, 22. This channel VI is called the primary channel and the gas flow Fl which flows through it is called the primary flow.

[0006] The rotor of the high-pressure compressor 16 is connected to the rotor of the high-pressure turbine 20 by a high-pressure shaft (not shown). The rotor of the low-pressure compressor 14 is connected to the rotor of the low-pressure turbine 22 by a low-pressure shaft (not shown) which passes through the high-pressure shaft and drives, directly or via a reduction gear, a blower propeller SI located generally upstream of the gas generator 12.

[0007] The blower propeller SI can be shrouded or unshrouded. In the case where it is shrouded as in the example shown, it is surrounded by a nacelle 30 which is connected to the gas generator 12 by radial arms 32.

[0008] The blower propeller SI generates an airflow F, part of which enters the gas generator 12 to form the primary flow Fl, and part flows around the gas generator 12 to form a secondary flow F2 in a secondary vein V2 which is crossed by the radial arms 32.

[0009] A turbomachine 10 is thus formed of several rotating elements called rotors and several stationary elements called stators. It is important to ensure a seal between a rotor and a stator in order to prevent gas leaks and optimize the performance of the turbomachine. Gas leaks represent a flow that does not contribute to the operation of the turbomachine and which, in fact, reduces its efficiency.

[0010] There are several types of sealing assemblies and the present invention relates more particularly to a labyrinth sealing assembly, an example of which is illustrated in [Fig.2].

[0011] A labyrinth seal assembly 40 comprises a rotor 42 which is movable in rotation about an axis, which is generally the axis A of the turbomachine 10 ([Fig.1]), and which is surrounded by a stator 44.

[0012] The rotor 42 has rotor blades, generally annular, 46 which extend radially outwards with respect to the axis A and which are surrounded by an abradable element, generally annular, 48 of the stator.

[0013] The slats 46 are able to cooperate, in particular by friction, with the element 48 to limit the passage of a gas flow F3 in operation from upstream to downstream between the slats 46 and the element 48, this gas flow F3 being a leakage flow and therefore having to be as low or small as possible.

[0014] The low-pressure turbine 22 of [Fig. 1] may, for example, include sealing assemblies 40 of this type. Each turbine stage comprises a bladed runner and a bladed distributor. The bladed distributor is attached to a housing that carries an abradable element 48. The distributor, the housing, and the abradable element 48 form part of the turbine stator. The bladed runner forms part of the turbine rotor and includes, on its outer periphery, blades 46 that are surrounded by the abradable element 48. The blades 46 of the bladed runner and the abradable element 48 of the stator form a labyrinth seal assembly 40. [Fig. 2] illustrates this configuration.

[0015] In another scenario, the bladed distributor carries on its inner periphery an abradable element 48 and the bladed wheel is connected to a rotor drum which carries rotor blades 46 surrounded by the abradable element 48. The blades 46 and the abradable element 48 form a labyrinth sealing assembly 40.

[0016] In a labyrinth seal assembly, pressure losses are mainly due to singular pressure losses (successive sudden expansions and contractions) but also to regular pressure losses due to gas friction on the parts (mainly on the rotor due to the rotation speed).

[0017] These frictions on the rotor are maximized when the tangential velocity differential between, on the one hand, the gas flow F3 entering the labyrinth and, on the other hand, the rotor wall is maximum.

[0018] Now, in a standard way, the gas flow F3 supplying a sealing assembly 40 has a velocity y with a strong tangential component v0, close to that Z7 of the rotor, which does not favor friction with the rotor (cf. [Fig.3]) because, in the rotor's frame of reference, the gas flow F3 will not be seen moving with a large relative tangential velocity.

[0019] The problem we seek to solve is to reduce the permeability of a labyrinth seal assembly to improve the efficiency of a turbomachine.

[0020] In the document FR-A1-3 099 788 cited above, it is proposed to place flow straighteners between the flaps of a labyrinth assembly, the purpose of which is to straighten the airflow exiting the assembly, to give the gas flow a non-zero gyration close to that of the gas flow in the turbine stream. This makes it possible to minimize the circumferential velocity differences between the gas flow ejected from the labyrinth assembly and the gas flow in the turbine stream, and therefore to minimize the so-called "mixing" losses of the turbine.

[0021] In the document FR-A1-3 071 540 cited above, a circumferential wall (solid therefore in the tangential direction) is proposed to create a separation of the boundary layer of the gas flow and hinder the entry of this gas flow into the labyrinth.

[0022] The present invention provides a solution to the problem mentioned above, which is simple, effective and economical.

[0023] The present invention proposes a solution to at least some of the problems of the prior art, which is simple, effective and economical. Summary of the invention

[0024] The invention relates to a labyrinth seal assembly for an aircraft turbomachine centered on an axis and extending circumferentially around this axis, comprising a rotor that is rotatable about the axis and surrounded by a stator, the stator comprising an abradable element that surrounds the rotor and delimits an annular passage around the rotor, the rotor comprising rotor blades that extend radially outwards with respect to the axis, these rotor blades passing through said passage and being capable of cooperating with the abradable element to limit the flow of an operating gas stream from upstream to downstream in said passage, characterized in that it further comprises, at the level of said passage and upstream of all the rotor blades, stator fins that project radially inwards from the abradable element, the stator fins extending axially opposite the rotor blades.

[0025] The invention makes it possible to improve the tangential velocity differential parameter between the gas flow and the rotor at the inlet of the maze. What distinguishes the invention from prior solutions is, in particular, the fact that vanes are arranged upstream. The blades are also present. Another difference concerns the function of these blades, as they straighten the gas flow, advantageously limiting or even stopping any gyration of the gas flow upstream of the labyrinth. Thus, the gas flow can enter the labyrinth with a predominantly axial component, preferably a purely axial one, or even with a slight tangential component opposite to the direction of rotor rotation. This increases the difference in tangential velocities between the incoming gas flow and the rotor of the labyrinth assembly, maximizes pressure losses, and minimizes the permeability of the assembly.

[0026] The sealing assembly according to the invention may comprise one or more of the following features, taken individually or in combination with each other:

[0027] — the stator fins are configured so as to axially straighten said flow of gas downstream of the stator fins;

[0028] — the stator fins are profiled or of constant thickness;

[0029] — each of the fins has a straight shape;

[0030] — the abradable element is annular;

[0031] — the rotor wipers are annular;

[0032] — each of the fins has a general streamlined, non-curved and symmetrical shape;

[0033] — each of the fins has a generally straight, non-curved shape and a thickness constant; • each of the stator fins has a generally curved shape; • Each of the stator fins has an aerodynamic profile and includes an intrados and an extrados connected together upstream to form a leading edge of the gas flow, and downstream to form a trailing edge of this gas flow; • each of the stator fins has an entry angle between 30° and 90°, this entry angle corresponding to the orientation angle of a part of the fin including the leading edge and being measured with respect to the axis and in a plane tangent to a circumference centered on the axis and passing through the fin;

[0034] — the entry angle is measured in a plane tangent to a circumference centered on the axis and passing through the blade; alternatively, this plane has a normal radial direction of the blade; • each of the stator fins has an exit angle between -20° and +10°, and preferably between -10° and 0°, this exit angle corresponding to the orientation angle of a part of the fin including the trailing edge and being measured with respect to the axis;

[0035] — the exit angle is measured in a plane tangent to a circumference centered on the axis and passing through the blade; alternatively, this plane has a normal radial direction of the blade; • each of the stator fins has a maximum radial height or dimension which represents between 20 and 50% of a maximum radial height or dimension of the most upstream fin; • the maximum radial height or dimension of each of the stator fins represents between 30 and 40%, and for example 33%+ / -2%, of the maximum radial height or dimension of the most upstream blade; • the maximum height or radial dimension of each of the stator fins is between 0.2cm and 1cm, preferably between 0.3cm and 0.8cm, and more preferably between 0.4cm and 0.6cm; • each of the stator fins has a height or radial dimension that decreases from upstream to downstream; • the stator fins are located at a minimum axial distance from the most upstream scraper which is less than or equal to a minimum axial inter-scraper distance; • each of the stator fins has an axial length or dimension that represents 20% to 90% of an axial length or dimension measured from an upstream end of the abradable coating to an upstream end of the most upstream slick;

[0036] — each of the fins has a maximum thickness / axial chord ratio approximately equal to 5.5 and preferably between 4 and 7; the maximum thickness is defined by the largest diameter of a circle that can be inscribed within the fin profile; the axial chord is defined as the length of the profile projected onto the aforementioned axis; • each of the stator fins has a maximum radial height or dimension which represents between 30 and 80% of a radial height or dimension of said passage which is measured at the level of the fins;

[0037] — the stator fins are regularly distributed around the axis;

[0038] — the number of fins is greater than 50 and preferably greater than 100;

[0039] — there are 2 to 5 licks.

[0040] The present invention also relates to a turbine or turbomachine for an aircraft, comprising at least one sealing assembly as described above. Brief description of the figures

[0041] Other features and advantages of the invention will become apparent upon reading the detailed description that follows, for an understanding of which reference should be made to the accompanying drawings in which:

[0042] [Fig-1] [Fig.1] is a schematic axial cross-sectional view of a turbomachine aircraft;

[0043] [Fig.2] [Fig.2] is a partial schematic cross-sectional view of an assembly labyrinth sealing;

[0044] [Fig.3] [Fig.3] is a very schematic partial view of a sealing assembly labyrinth such as that of [Fig.2];

[0045] [Fig.4] [Fig.4] is a very schematic partial view of a sealing assembly labyrinth according to the invention;

[0046] [Fig. 5] [Fig. 5] is a schematic cross-sectional view of an embodiment of a fin for a sealing assembly according to the invention;

[0047] [Fig. 6] [Fig. 6] is a schematic cross-sectional view of an alternative embodiment of a fin for a sealing assembly according to the invention;

[0048] [Fig.7] [Fig.7] is a partial schematic cross-sectional view of an embodiment of a sealing assembly according to the invention;

[0049] [Fig.8] [Fig.8] is a partial schematic cross-sectional view of a variant of fabrication of a sealing assembly according to the invention;

[0050] [Fig. 9 Figure 9 is a partial schematic perspective view of an assembly of sealing and illustrates the operation of the assembly.

[0051] [Fig. 10] [Fig. 10] is a schematic cross-sectional view of an alternative embodiment of a fin for a sealing device according to the invention;

[0052] [Fig. 1la-11b] Figures 1la and 11b are schematic cross-sectional views of a variant of the embodiment of a fin for a sealing device according to the invention. Detailed description of the invention

[0053] Figures 1 to 3 have been described above. These figures can be considered as illustrating the environment of the invention.

[0054] As mentioned above and illustrated in [Fig.2] in particular, the invention relates to a labyrinth seal assembly 40 for an aircraft turbomachine 10 such as that illustrated in [Fig.1].

[0055] The assembly 40 comprises a rotor 42 which is mobile in rotation around an axis A and which is surrounded by a stator 44.

[0056] The stator 44 includes an element, generally annular, abradable 48 which surrounds the rotor 42 and which delimits around the rotor 42 an annular passage P.

[0057] The rotor 42 has blades, generally annular, of rotor 46 which extend radially outwards with respect to the axis A.

[0058] These slats 46 cross the passage P and are able to cooperate with the abradable element 48 to limit the flow of a gas flow F3 in operation from upstream to downstream in the passage P.

[0059] Fig. 4 illustrates the principle of the invention, which consists of providing fins 50 at the level of the passage P and upstream of all the blades 46 of the rotor 42.

[0060] The fins 50 are supported by the stator 44 and are preferably configured to straighten the gas flow F3. The fins 50 can be profiled or of constant thickness.

[0061] Advantageously, the fins 50 make it possible to cancel the tangential component of the gas flow F3 at the inlet of the labyrinth seal or to slightly reverse the tangential direction with respect to the direction of rotation of the rotor in order to increase the air friction on the rotor by increasing the differential of tangential velocities between the rotor walls and the incoming gas flow. This friction creates regular pressure losses and thus reduces the permeability of the labyrinth seal assembly 40.

[0062] The fins 50 thus preferably have the function of straightening the gas flow F3 to give it a predominantly axial component y? at the outlet of the fins 50.

[0063] Figures 5 and 6 show examples of aerodynamic profiles for the fins 50. In these figures, arrow F3 represents the orientation of the aforementioned gas flow and arrow F4 represents the direction of rotation of the rotor 42.

[0064] Each of the fins 50 has, for example, a generally curved shape.

[0065] Each of the fins 50 preferably has an aerodynamic profile and includes a intrados 50a and an extrados 50b connected together upstream to form an edge 50c of attack of the gas flow F3, and downstream to form an edge 50d of escape of this gas flow F3.

[0066] Each of the fins 50 can also have a straight and rectilinear shape, profiled or not.

[0067] Each of the fins 50 preferably has an entry angle α between 30° and 90°.

[0068] This entry angle corresponds to the orientation angle of a part of the fin 50 comprising the leading edge 50c and being measured with respect to the axis A and preferably in a Z plane tangent to a circumference centered on the axis A and passing through the fin 50. The Z plane can be considered as the drawing plane.

[0069] Each of the fins 50 preferably has an exit angle [3] between -20° and +10°, and preferably between -10° and 0°.

[0070] This exit angle [3 corresponds to the orientation angle of a part of the fin 50 comprising the trailing edge 50d and being measured with respect to the axis A and preferably in the aforementioned Z plane.

[0071] In the embodiment of [Fig.5], each of the fins 50 has an entry angle a of the order of 30° and an exit angle [3 of the order of 0°.

[0072] In the embodiment of [Fig.6], each of the fins 50 has an entry angle a of the order of 90° and an exit angle [3 of the order of -10°. This angle is negative because it is opposite to the direction of rotation of the rotor (arrow F4).

[0073] Fig. 7 represents an embodiment or a first configuration of a labyrinth seal assembly 40 including the invention and consequently the fins 50.

[0074] In this configuration, the stator 44 of the assembly 40 is formed by a turbine distributor which carries on its internal periphery the abradable element 48.

[0075] In the example shown, the abradable element 48 is stepped and comprises two successive axial sections of different internal diameters. The upstream section has a smaller internal diameter than the downstream section.

[0076] The blades 46 are carried by a turbine drum which generally connects two turbine wheels. There are two blades 46, and they are respectively associated with the two sections of the lining 48.

[0077] The fins 50 are located at the upstream end of the coating 48 in the example shown. The upstream ends of the fins 50 can be aligned axially with the upstream end of the coating 48.

[0078] Moreover, each of the fins 50 has a radial height or dimension Ha which decreases from upstream to downstream in the example shown.

[0079] Fig. 8 represents another embodiment or a second configuration of a labyrinth seal assembly 40 including the invention and the fins 50.

[0080] In the variants of figures 10 and 11a, each of the profiled fins respects a maximum thickness emaxi / axial chord ratio approximately equal to 5.5 and preferably between 4 and 7. The maximum thickness emaxi is defined by the largest diameter of a circle that can be inscribed within the profile, and the axial chord Cax is defined as the length of the profile projected onto the axis A.

[0081] The fin 50 of [Fig. 10] is profiled, whereas the fins 50 of Figures 11a and 11b are straight. Figure 11a shows a fin 50 with a generally profiled, non-curved, and symmetrical shape. Figure 11b shows a fin 50 with a generally straight, non-curved shape and constant thickness.

[0082] In this configuration, the stator 44 of the assembly 40 is formed by a turbine ring sector which carries on its inner periphery the abradable element 48.

[0083] In the example shown, the abradable element 48 is stepped as mentioned above.

[0084] The blades 46 are carried by a turbine wheel. There are two blades 46, and they are respectively associated with the two sections of the coating 48.

[0085] Moreover, each of the fins 50 has a radial height or dimension Ha which decreases from upstream to downstream in the example shown.

[0086] In both of these configurations and generally speaking, each of the fins 50 can:

[0087] - have a maximum radial height or dimension Ha which represents between 20 and 50%, preferably between 30 and 40%, and for example 33% + / - 2% of a maximum radial height or dimension Ho of the upstreammost slat 46, and / or

[0088] - have a maximum radial height or dimension Ha which is between 0.2cm and 1 cm, preferably between 0.3 cm and 0.8 cm, and more preferably between 0.4 cm and 0.6 cm, and / or

[0089] - have a maximum radial height or dimension Ha which represents between 30 and 80% of a radial height or dimension Hx of the passage P which is measured at the level of the fins 50, and / or

[0090] - be located at a minimum axial distance La from the upstream lick 46 which is less than or equal to a minimum axial distance Lo between licks, that is, between two consecutive licks 46, and / or

[0091] - have an axial length or dimension Le which represents 20% to 90% of a axial length or dimension Lb measured from an upstream end of the abradable coating 48 to an upstream end of the most upstream lick 46.

[0092] Reference is now made to [Fig.9] which illustrates the operating principle of a sealing assembly 50 according to the invention.

[0093] In this figure:

[0094] - the arrow Fl represents the primary flow in the turbine stream,

[0095] - Arrow F3 represents the gas flow upstream of passage P of the labyrinth,

[0096] - arrow F4 represents the direction of rotation of the rotor 42 around the aforementioned axis, and

[0097] - the arrow F5 represents the gas flow exiting the fins 50 and therefore in the passage P of the labyrinth.

[0098] It is observed that the fins 50 make it possible to straighten the gas flow F3 and preferably to remove any tangential component so that the flow of the flux F5 in the passage P and in particular through the slats 46 has essentially an axial component.

[0099] This increases the friction of the gas flow on the rotor 42 by increasing the tangential velocity differential between the rotor and the gas flow. This friction creates pressure losses that decrease the permeability of the assembly 40.

Claims

Demands

1. A labyrinth seal assembly (40) for an aircraft turbomachine (10) centered on an axis (A) and extending circumferentially around this axis (A), comprising a rotor (42) that is rotatable about the axis (A) and surrounded by a stator (44), the stator (44) comprising an abradable element (48) that surrounds the rotor (42) and delimits around the rotor (42) an annular passage (P), the rotor (42) comprising rotor blades (46) that extend radially outwards with respect to the axis (A), these rotor blades (46) passing through said passage (P) and being able to cooperate with the abradable element (48) to limit the flow of an operating gas stream (F3) from upstream to downstream in said passage (P), characterized in that it comprises in addition, at the level of said passage (P) and upstream of all rotor scrapers (46),stator fins (50) which project radially inwards relative to the abradable element (48), the stator fins (50) extending axially opposite the rotor blades (46).

2. Sealing assembly (40) according to claim 1, wherein each of the stator fins (50) has a generally curved shape.

3. Sealing assembly (40) according to claim 1 or 2, wherein each of the stator fins (50) has an aerodynamic profile and comprises an intrados (50a) and an extrados (50b) connected together upstream to form a leading edge (50c) of the gas flow (F3), and downstream to form a trailing edge (50d) of this gas flow (F3).

4. Sealing assembly (40) according to claim 3, wherein each of the stator fins (50) has an entry angle (a) between 30° and 90°, this entry angle (a) corresponding to the orientation angle of a portion of the fin (50) comprising the leading edge (50c) and being measured with respect to the axis (A).

5. Sealing assembly (40) according to claim 3 or 4, wherein each of the stator fins (50) has an exit angle (|3) between -20° and +10°, and preferably between -10° and 0°, this exit angle (|3) corresponding to the orientation angle of a portion of the fin (50) comprising the trailing edge (50d) and being measured with respect to the axis (A) and in a plane (Z) tangent to a circumference centered on the axis (A) and passing through the fin (50).

6. Sealing assembly (40) according to any one of the preceding claims, wherein each of the stator fins (50) has a maximum radial height or dimension (Ha) which is between 20 and 50% of a maximum radial height or dimension (Ho) of the upstream fin (46).

7. Sealing assembly (40) according to claim 6, wherein the maximum radial height or dimension (Ha) of each of the stator fins (50) represents between 30 and 40%, and for example 33%+ / -2%, of the maximum radial height or dimension (Ho) of the upstream fin (46).

8. Sealing assembly (40) according to claim 5 or 6, wherein the maximum height or radial dimension (Ha) of each of the stator fins (50) is between 0.2cm and 1cm, preferably between 0.3cm and 0.8cm, and more preferably between 0.4cm and 0.6cm.

9. Sealing assembly (40) according to any one of claims 5 to 8, wherein each of the stator fins (50) has a radial height or dimension (Ha) which decreases from upstream to downstream.

10. Sealing assembly (40) according to any one of the preceding claims, wherein the stator fins (50) are located at a minimum axial distance (La) from the upstream fin (46) which is less than or equal to a minimum axial inter-fin distance (Lo).

11. Sealing assembly (40) according to any one of the preceding claims wherein each of the stator fins (50) has an axial length or dimension (Le) which represents 20% to 90% of an axial length or dimension (Lb) measured from an upstream end of the abradable coating (48) to an upstream end of the most upstream nozzle (46).

12. Sealing assembly (40) according to any one of the preceding claims, wherein each of the stator fins (50) has a maximum radial height or dimension (Ha) which represents between 30 and 80% of a radial height or dimension (Hx) of said passage which is measured at the fins (50).

13. Turbomachine (10) for an aircraft, comprising a turbine which has at least one sealing assembly (40) according to any one of the preceding claims.

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